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TECHNICAL MEMORANDUM

X-215

EFFECT OF AFTERBODY TERMINAL FAIRING8 ON THE PERFORMANCE

OF A PYLON-MOUNTED TURBOJET-NACELLE MODEL

By Conrad M. Willis and Charles E. Mercer

Langley Research Center

Langley Field, Va.

Declassified March 15, 1962

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON March 1960

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

TECHNICAL MEMORANDUM X-215

EFFECT OF AFTERBODY TERMINAL FAIRINGS ON THE PERFORMANCE

OF A PYLON-MOUNTED TURBOJET-NACELLE MODEL*

By Conrad M. Willis and Charles E. Mercer

SUMMARY

An investigation of the effect of afterbody terminal fairings on

the performance of a pylon-mounted turbojet-nacelle model has been con-

ducted in the Langley 16-foot transonic tunnel. A basic afterbody havinga boattail angle of 16 ° was investigated with and without terminal

fairings. The equivalent boattail angle, based on the cross-sectional

area of the afterbody and terminal fairings, was 8°. Therefore, a simple

body of revolution with a boattail angle of 8° was included for compari-

son. The tests were made at an angle of attack of 0°, Mach numbers

of 0.80 to 1.05, jet total-pressure ratio of 1 to approximately 5, and

an average Reynolds number per foot of 4.1 x 106 • A hydrogen peroxide

Jet simulator was used to supply the hot-Jet exhaust.

The results indicate that addition of terminal fairings to a

16 ° boattail afterbody increased the thrust-minus-drag coefficients and

provided the lowest effective drag of the three configurations tested.

INTRODUCTION

0ptimumperformance of a nozzle-exit--afterbody combination at both

subsonic and supersonic speeds requires continuously variable internal

and external surfaces. Such variable geometry configurations result in

complexity of fabrication and weight penalty. In previous attempts to

bircumvent these problems, simple semifixed geometry configurations have

been designed that perform well at high speeds. Predominant among these

configurations has been the fixed convergent-divergent ejector with

variable primary-nozzle and secondary air flow (refs. 1 to 3). However,

these configurations show sizable performance losses when operated at

off-design conditions. The use of terminal fairings as a new approachto the solution of this problem was introduced in reference 4. The

Title, Unclassified.

2

basis for this concept of design is the interaction of the internal andexternal flows in the afterbody Jet-exit region. The results of a pre-vious investigation related to this subject have been reported in refer-ence 5. A terminal-fairing configuration consists of a multiplicity ofstreamlined bodies clustered around the afterbody and extending down-stream of the jet exit and spaced so as not to form a complete barrierbetween jet and external flow.

Because of the complex nature of the _xing flows around the ter-minal bodies or fairlngs, changes in afterbody boattailing and thearrangement and shape of the fairings mayhave appreciable effects onthe performance of nacelles with this type of afterbody-nozzle combi-nation. The terminal fairings of reference 5 consisted of six bodiesof circular cross section on an afterbody with a boattall angle ofabout 16°. The present paper reports results of a continuation of theinvestigation of terminal fairings; however, the terminal fairings con-sisted of only four bodies of flattened cross-sectional shape. The per-formance of a simple body with 16° boattaiL angle is comparedwith thatof the sameafterbody having the four ter_nal fairings added. Theseadded bodies produced a configuration havimg an area equivalent to abody of revolution with an 8° boattail angle (a near-optimum value).A simple body of revolution having a boattE_il angle of 8° was also testedto provide a further comparison. All conf2gurations were tested with anonafterburning type primary nozzle.

This investigation was conducted in the Langley 16-foot transonictunnel over a Machnumberrange of 0.80 to 1.05 at 0° angle of attack.Jet total-pressure ratio was varied from 1 (Jet off) to approximately 5at each Machnumber. The effects of secondary air on the terminal-fairing configuration and a comparable bo_r of revolution were investi-gated at a Machnumber of 0.90 and a Jet t,_tal-pressure ratio of 4; thecorrected secondary-to-primary weight-flow ratios varied from 0 (no flow)to about 0.06. The turbojet exhaust was simulated by a hydrogen perox-ide hot-jet unit similar to that described in reference 6.

SYMBOLS

A

CD

CD'

CD,a

cross-sectional area, sq ft

drag coefficient

effective drag coefficient, (c;-co)

afterbody pressure-drag coefficient_ -CpAz

Amax/,

5

CF

CF - CD

CF,ej

CFi,c

CF,p

Cp

D

d

F

F - D

thrust coefficient

F- D

thrust-minus-drag coefficient, qAn_ x

ejector jet thrust coefficient,FeJ

qAm_x

ideal convergent-nozzle Jet thrust coefficient,

priory-nozzle jet thrust coefficient, FpqAm x

PZ - P_

pressure coefficient, q

drag, Ib

diameter, in.

thrust, ib

thrust minus drag, Fba I - Dba I + (As, 2 - As,I)(Pl - P2), ib

Fej

Fi,c

Fp

g

L

M

ejector jet thrust,

A e

F_ + v3 + (p3 - _)(As - A_) ÷g _As

ideal convergent-nozzle jet thrust,

-g-Wp_gR _2 Tt J + Ap(pp - p_), iby+l '

primary-nozzle thrust, ib

acceleration due to gravity, ft/sec 2

afterbody length, in.

distance from primary nozzle to afterbody exit, in.

free-streamMach number

4

P

Pt

Pt, J/P_

q

R

T t

V

w

x

Y

7

B

8

Subscripts:

bal

e

eqv

J

n_x

P

static pressure, lb/sq ft

total pressure, lb/sq ft

Jet-pressure ratio (ratio of primary Jet total pressure to

free-stream static pressure)

dynamic pressure, lb/sq ft

gas constant, ft/°R

stagnation temperature, OF

velocity, ft/see

weight flow rate, lb/sec

corrected secondary-to-primary weight-flow ratio

axial distance from reference _tations (see figs. 1 and 5),

positive rearward, in.

radial coordinate, in.

ratio of specific heats

boattail angle of afterbody ba_e, deg

meridian angle of model, deg

balance

exit of afterbody

equivalent

jet

local

maximum

primary nozzle

s

OO

1

2

3

seal or secondary air

free-stream conditions

forward compartment of model

outer compartment of model

rear compartment of model

APPARATUS AND PROCEDURE

Wind Tunnel and Model Support System

The investigation was conducted in the Langley 16-foot transonic

tunnel, which is an octagonal slotted-throat single-return wind tunnel

operated at atmospheric stagnation pressures. The model was supported

by a sweptback pylon attached to a conventional sting 18 inches below

the model center line as shown in figure i. Since the model pylon is

similar to actual installations and since the same support was used for

all configurations, no corrections were made for support interference.

Interference effects for this mounting system are discussed in

reference 7.

Nacelle and Balance System

A sketch of the nacelle model is presented in figure I, and a

photograph of the nacelle and pylon is shown in figure 2. The nacelle

shell and jet simulator unit were separate systems and each was attached

to the pylon by its own balance. The hydrogen peroxide Jet simulator

(described in ref. 6) had an exhaust temperature of about 1,350 ° F.

Secondary air was exhausted into an annular passage between the tail-

pipe and nacelle shell.

Configurations

The three afterbody configurations (fig. 3) were designed for the

purpose of evaluating the relative performance of: a basic axisymmetric

boattailed afterbody (configuration I), the same afterbody with terminal

fairings added to reduce the effective boattail angle (configuration II),

and a simple afterbody having axisymmetric boattailing equivalent to that

determined by the axial distribution of cross-sectional area of the

terminal-fairing configuration (configuration III). All these after-

bodies had diameter ratios (jet nozzle to maximum nacelle and base to

6

maximum nacelle) that corresponded to those for typical turbojet-nacelle

installations with primary nozzles in the r_onafterburning condition.

The afterbodies were detachable at the 47.125-inch station.

Dimensions of configuration I, the basic 16 ° boattail, are shown

in the sketch presented as figure 3(a). Configuration II (fig. 3(b))

was formed by adding four detachable terminal fairings to configuration I.

The fairings were of flattened cross-sectional shape and were designed to

provide an equivalent boattail angle of 8°. This angle was arbitrarily

defined as the boattail angle produced by distributing the cross-

sectional area of the four fairings in an annulus around the basic boat-

tail and was measured at the 57.030-inch afterbody exit station. Con-

figuration III (fig. 3(c)) is representative of a low-drag afterbody in

the transonic speed range (ref. 8). The low boattail angle necessitated

an extension in afterbody length to achieve a base area approximately

the same as the other two configurations (figs. 4 and 5). Configura-

tion III was selected for testing to provide a performance comparison

between the terminal-fairing configuration _nd a simple body of revolu-

tion with the same boattail angle. Area distributions for configura-

tions I, II, and III are shown in figure 5.

Instrumentation

External and internal static pressures were measured on the after-

bodies at locations shown in figture 3. It should be noted that for con-

figurations I and II there is only one row _f external pressure orifices

which is on the top of the afterbody. In aldition, primary jet total

pressures, secondary air exit static pressures, and primary and second-

ary total temperatures were measured. (See fig. 6.) The pressure

tubing from each orifice was conducted out _f the nacelle through the

pylon support and connected to an electrical pressure transducer located

in the sting barrel. The electrical pressure transducers were manifolded

to a common reference pressure and the whol_ transducer manifold system

was held at a constant temperature to keep aoth the zero and sensitivity

shifts of the transducers to a minimum. El_ctrlcal signals from the

pressure transducers were transmitted to cacrier amplifiers and then to

recording oscillographs located in the tunn_l control room.

The thrust forces of the jet simulator were obtained from a one-

component thrust balance. A four-component internal balance measured

the forces and moments on the nacelle; howe_er, only the drag measure-ments are presented in this paper. Figure ) indicates the balance loca-

tions and the pressures, areas, and tempera;_es associated with the

reduction and correction of data. An elect:_onic flowmeter and a cali-

brated venturi were used to measure the primary and secondary flow rates,respectively.

7

Data Reduction

Model data recorded by oscillograph trace deflections were used to

compute standard force and pressure coefficients. Because of limited

instrumentation of configurations I and II, afterbody pressure drag was

not computed for these configurations.

Since thrust and drag cannot be readily separated for configurations

designed to allow mixing of internal and external flows, thrust minus

drag, or net propulsive force, is used to compare the three afterbodies.

The processes described in reference 5 were used to obtain net propul-

sive force and effective drag. Ejector thrust was determined as follows

for configuration III:

Ws ]A AeFej=Fp+_ Vs+(P3-p_)(As-Ap)+ (P_-P_)_s

where

Ws V3Fp=Fbal-T -(P3-P_)(As,2-Ap)+(Pl-P_)As,2

The equation for Fp applies to all configurations. Locations of these

pressures and areas are shown in figure 6.

Accuracy

Estimated accuracy of data presented in this paper is as follows:

M ............................... ±0.005

Pt,j/p _ ............................ ±O.lO

CD, a ............................. ±0.01

CF,ej ............................. ±0.01

CF - CD ............................ ±0.01

CD' .............................. ±0.01

Ws_ ........................... ±0.oo5

Wp _ Tt,p

TESTS

All tests were conducted at 0° angle of attack. The Mach number

range was from 0.80 to 1.05, and the average Reynolds number per foot

8

was 4.1 × 106. Ratios of primary Jet total pressure to free-stream

static pressure ranged from 1 (jet off) to about 5 at each Mach number.

Secondary air at flow rates of 0 to about 0_25 pound per second

J(Ws_ITt's = 0 to 0.06)" was used for tests of configurations II and III VTt,p\ !

at a Mach number of 0.90 and a Jet total-pressure ratio of 4.

RESULTS

Longitudinal distributions of pressure for afterbody configura-

tion III at pressure ratios of 1 and 5 are presented in figure 7. In

figure 8, representative pressure distributions for the three afterbody

configurations are compared. Figure 9 presents pressure distributions

obtained over the six-body terminal-fairing configuration of reference 5

and compares these distributions with those obtained over the four-body

terminal-fairing configuration of the present investigation. Afterbody

pressure-drag coefficient for configuration IIl is presented in figure lO.

Thrust data are shown in figures ll to 13. Figures 14 and 15 present

performance comparisons on a thrust-minus-drag basis. Effective drag

coefficient at a scheduled Jet total-pressuze ratio is shown in figure 16.

Thrust-minus-drag coefficients are compared for various configurations

in figures 17 and 18.

DISCUSSION

Afterbody Pressure Distributions

The effect of Jet operation on afterbody pressure distributions of

configuration III is presented in figure 7. Jet operation increased

afterbody pressures near the base. This favorable Jet interference gen-

erally decreased with increasing Mach number, an effect that is typical

for boattails of this shape (ref. 7).

Figure 8 presents a comparison of pressure distributions obtained

from the top row of orifices for the three a_terbody configurations of

the present investigation. Configuration I _ith a boattail angle of 16 °

had the most negative afterbody pressures, as expected. The pressure

level for the terminal-fairing configuration with an equivalent boattail

angle of 8° (configuration II) generally fell about halfway between the

levels for the bodies of revolution with boa_tail angles of 16 ° and 8°

(configurations I and III). However, in the region near the boattail

base where the orifices were located between the terminal fairings, the

pressures for configuration II were more negative than those for either

9

of the bodies of revolution. Jet operation decreased the pressures near

the base for configuration II but increased these pressures for theother two afterbodies.

In order to examine this apparently unfavorable effect of terminal

fairings on the boattail base region, sample pressure distributions

obtained over the six-body terminal-fairing configuration of reference 5

are shown in figure 9. The data for configuration II shown in figure 8

at a Mach number of 0.90 are repeated for comparison purposes. Although

the pressure coefficients for the six-body terminal-fairing configura-

tion were more positive than those for the four-body terminal-fairing

configuration_ the trends of the pressure distributions near the bases

of the bodies were similar. The pressures on the surfaces of the six

terminal fairings behind the base generally increased substantially

with jet operation and these increased pressures resulted in thrust

forces on the fairings. It would be expected that a similar pressure

recovery would occur over the four terminal bodies of the present

investigation.

Afterbody Pressure Drag

Afterbody pressure-drag coefficients for configuration III are

shown in figure lO. Increasing the jet total-pressure ratio caused

decreases in afterbody pressure-drag coefficient. Limited data showing

the addition of secondary air flow at M = 0.90 indicated little

effect on afterbody drag in this investigation. The decrease in after-

body pressure-drag coefficient indicated by the test point at the Jet-

off condition is probably due to a base bleed effect. Also shown in

figure lO are data for an afterbody with a boattail angle of 15 ° and a

base-to-maximum-diameter ratio of 0.538 (afterbody II of ref. 7). It

would be expected that the magnitude of the afterbody pressure-drag

coefficients for the 15° boattail body of reference 7 would be approxi-

mately the same as those for the 16° boattail afterbody of the present

investigation (configuration I) since both configurations were investi-

gated on the same nacelle and support system. The difference in level

of afterbody pressure-drag coefficients for the two configurations pre-

sented in figure lO should therefore be indicative of the drag differ-

ences expected between configurations I and III of the present

investigation.

Primary-Nozzle Jet Performance

The variation of primary-nozzle jet thrust coefficient with Jet

total-pressure ratio is presented in figure ii. The data are compared

with the ideal convergent-nozzle jet thrust coefficient based on meas-

ured Jet total pressure, temperature, and weight flow rates. Since the

i0

sameprimary nozzle was used for all confi_ations, the test data shouldfall on a single line. Efficiency, as indi(:ated by the ratio of primarythrust to ideal thrust, varied from approximately 0.90 at Pt,j/p _ = 2to 0.95 at Pt,jJP_ = 5.

Ejector Thrust

In order to obtain a low boattail angle with the samebase diameteras the other configurations, the afterbody of configuration III had tobe extended and thereby resulted in an ejector with a greater spacingratio. (See fig. 3(c).) This arrangement, therefore, would require acarefully programed amount of secondary air flow if it were to operateefficiently as an ejector. Reference 9 indicates a thrust ratio ofapproximately 1.0 at zero secondary air flows for ejector geometriessimilar to configurations I and II in the pressure-ratio range of thisinvestigation. Ejector thrust coefficient :for configuration III is pre-sented in figure 12 for zero secondary air flow. In addition, at a Machnumberof 0.90, a point is presented for th_ maximumamount of secondaryair flow available through the system. Wit]l no secondary air flow,large ejector-thrust losses occur at the higher Jet total-pressure ratios.These losses are probably due to jet attachment to the shroud and lowpressures in the secondary air passages. With the addition of about6-percent corrected secondary air flow, the ejector thrust coefficientsapproached more closely the ideal convergent-nozzle thrust coefficient.

An indication of the ejector performance with secondary weight flowratio maybe noted from figure 13 where Jet thrust ratios of configura-tion III and static-test data for a similar ejector (diameter ratio, 1.40;spacing ratio, 0.803) of reference 9 are c_apared. It can be seen thatthe trends with jet total-pressure ratio ar_ similar for the two configu-rations at zero secondary air flow, but the losses for configuration IIIare muchhigher, probably due to the differences in the internal geometryof the secondary-flow passage and Machnumbereffects. With each suc-cessive increase in secondary air flow, an Lncrease in performance wasobtained with the test configuration. This increase indicated that, withsufficient secondary air flow_ the ejector _f configuration III wouldprovide acceptable performance. However, ti_ese gains would be offsetby the penalty for bringing this secondary _ir on board. The forcerequired to bring 6-percent corrected seconlary air flow to rest fromthe free-streamMach numberof 0.90 and Pt,j P_ = 4 amountsto a pen-alty of about 0.084 in drag coefficient.

Thrust-Minus-Drag Measurements

Thrust-minus-drag measurementsprovide a convenient meansof com-paring overall performance of configurations having the sameprimary

CONFIDENTIAL ii

nozzles. Separation of the data into the basic quantities of thrustand drag necessitates an arbitrary division of forces between the thrustand the external nacelle drag. This division becomesparticularly com-plicated for the terminal-fairing configuration because of the ejectoraction of the Jet bulb expanding along the inner surface and sides ofthe fairings.

The variation in thrust-minus-drag coefficient with Jet total-pressure ratio for the three afterbody configurations is shownin fig-ure 14. The addition of four terminal fairings to the 16° boattailbody improved the model performance at all Jet total-pressure ratios.Performance losses for configuration III, previously indicated by theejector-thrust-coefficient curves of figure 12, occurred at the higherjet total-pressure ratios. With approximately 6-percent corrected sec-ondary air flow at a Machnumberof 0.90, the performance of configura-tion III was slightly better than that of configuration I when lossesdue to obtaining this flow were neglected. The addition of secondaryair flow, however, had little effect on the performance of the terminal-fairing configurations.

Performance comparisons are presented in figure 15 for a typicalschedule of turbojet-engine pressure ratios with Machnumber. Gains ofabout 6 percent in thrust-mlnus-drag performance in the Machnumberrange from 0.90 to 1.O0 were obtained by adding the terminal fairingsto the basic body. It should be noted that thrust-minus-drag perform-ance for configuration III is penalized by the absence of secondary airflow.

The data of figure 15 are presented in another form in figure 16to show the variation with Machnumber of the effective drag coeffi-cients of the three configurations considered in this paper. Effectivedrag coefficients were obtained by subtracting the experimentally deter-mined values of thrust-minus-drag coefficient from the computedvaluesof primary-nozzle thrust coefficient. The data for configurations Iand II showthat the addition of the terminal fairings to the basic con-figuration reduced the effective drag 44 percent at a Machnumber of 0.90and about 21 percent at a Machnumberof 1.O0. Since the effective dragcoefficients reflect gains or losses associated with the internal ejec-tor arrangement as well as differences in external drag (see ref. 5),the losses in ejector thrust for configuration III (fig. 12) would showup as high effective drag coefficients. Therefore, data for configura-tion III without secondary flow are omitted from figure 16. A one-pointcomparison of the terminal-fairing and the 8° axisymmetrical boattailbodies (configurations II and III) is madeat a Machnumber of 0.90 toindicate the relative effective drag coefficients of the two afterbodieswith a representative corrected secondary-to-primary weight flow ratioof 0.04. With this secondary air flow rate, the four-terminal-fairingmodel had an effective drag coefficient about 0.047 lower than that for

12

the 8° axisymmetrical boattail model. These results were obtained withthe nonafterburning primary Jet nozzle, and performance improvementsdue to the terminal fairings would probably be larger for afterburnernozzle operation. (See ref. 5.) Thrust-minus-drag coefficients fortwo terminal-fairing configurations are presented in figure 17. Sincethe nozzle sizes and propellant flow rates were different, the datawere normalized on the basis of primary-nozzle exit area. The configu-ration with four terminal fairings provided better performance than thesix-terminal-fairing configuration at Machnumbersabove 0.95. However,all the differences shownin propulsive fcrce are not necessarily dueto the change in the numberof fairings, but must be attributed to theentire afterbody arrangement.

The effect of secondary air weight flow ratio on the thrust-minus-drag coefficients of several types of terminal-fairing configurationsand the comparative afterbody, configuration III, is shownin figure 18for a Machnumberof 0.90 and a Jet total-pressure ratio of 4. Theslotted divergent ejector included in this figure (ref. 5) is consideredto also represent a type of terminal fairing in that the internal por-tion of the body is ventilated to the free stream beyond the primarynozzle. In general, small amounts of seccndary air flow produced themost change in thrust-minus-drag coefficients. Additional increases insecondary air flow rates resulted in only small changes in thrust-mlnus-drag coefficients. The four-terminal-fairing configuration had higherthrust-minus-drag coefficients than the other afterbodies consideredherein throughout the secondary air flow range of this investigation.

CONCLUSIONS

An investigation of the effects of afterbody terminal fairings onthe performance of a pylon-mounted Jet-na(elle model has been conductedin the Langley 16-foot transonic tunnel. The results have led to thefollowing conclusions:

1. Addition of four terminal fairingE to a simple 16° boattail bodyof revolution increased the thrust-minus-drag coefficients and decreasedthe effective drag coefficients over the R_ch numberrange.

2. At Machnumbersabove 0.95, the configuration with four terminalbodies had higher thrust-minus-drag coefficients (based on primary-nozzleexit area) than did the configuration witk six terminal bodies of aprevious investigation (NASAMEMO10-24-5EL).

Langley Research Center,National Aeronautics and SpaceAdministration,

Langley Field, Va., October 15, 1959.

13

REFERENCES

i. Greathouse, William K., and Beale, William T.: Performance Charac-

teristics of Several Divergent-Shroud Aircraft Ejectors. NACA

RM E55G21a, 1955.

2. Beheim, Milton A.: Off-Design Performance of Divergent Ejectors.

NACA RM E58GlOa, 1958.

3. Norton, Harry T., Jr., Cassetti, Marlowe D., and Mercer, Charles E.:

Transonic Off-Design Performance of a Fixed Divergent Ejector

Designed for a Mach Number of 2.0. NASA TM X-165, 1959.

4. Swihart, John M., Norton, Harry T., Jr., and Schmeer, James W.:

Effect of Several Afterbody Modifications Including Terminal

Fairings on the Drag of a Single-Engine Fighter Model With Hot-

Jet Exhaust. NASA MEMO I0-29-58L, 1958.

5. Runckel, Jack F.: Preliminary Transonic Performance Results for

Solid and Slotted Turbojet Nacelle Afterbodies Incorporating Fixed

Divergent Jet Nozzles Designed for Supersonic Operation. NASA

MEMO I0-24-58L, 1958.

6. Runckel, Jack F., and Swihart, John M.: A Hydrogen Peroxide Hot-Jet

Simulator for Wind-Tunnel Tests of Turbojet-Exit Models. NASA

MEMO 1-10-59L, 1959.

7. Swihart, John M., Mercer, Charles E., and Norton, Harry T., Jr.:

Effect of Afterbody-Ejector Configurations on the Performance at

Transonic Speeds of a Pylon-Supported Nacelle Model Having a Hot-

Jet Exhaust. NASA MEMO I-4-59L, 1959.

8. Cubbage, James M., Jr.: Jet Effects on the Drag of Conical After-

bodies for Mach Numbers of 0.6 to 1.28. NACA RML57B21, 1957.

9. Greathouse, W. K., and Hollister, D. P.: Preliminary Air-Flow and

Thrust Calibrations of Several Conical Cooling-Air Ejectors With

a Primary to Secondary Temperature Ratio of 1.O. II - Diameter

Ratios of 1.06 and 1.40. NACA RME52F26, 1952.

14

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EXTERNAL COORDINATES

Sto x l(Radius]

47'.125 0.000 3,25048.260 1.135:5,23850.260 3.135 $.17552.260 5.155 5.05555.260 6.135 2.97554.460 7.535 2.86055.260 8.135 2.77256.260 9.135 2.65057.0:50 9.905 2.54657.260 10.135 2.514

58.260 11.135 2.57358.550 11.425 2.333

EJECTOR GEOMETRY ',

Exit to jet diameter ....... 1.437 i

ratio, de/d p

Spacing ratio, L/dp ...... 0.753

Row

e=6_ 57_106_ 159°and 315 °

ORIFICE LOCATIONS

EXTERNAL

x xl/.

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x x/Z,

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Figure 3.- Concluded.

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Jet off

J \

0

J

L_ r

J

- Pt,j/Poo =5

.I--<..-\cf x....

.t0 .2 .4 6

/f \F ',. \

-" " 2,-._. "<r__""u. ""zk

E_ :X

\

\

,/

X'_ "o.,%

- %

ff-- \'k

.8 1.0 0 .2 .4 .6 .8Fraction of afferbod/ length, x/Z,

\k

\\.

(a) t4 = 0.8o.

Figure 7.- Effect of jet operation on pressure distributions of

configuration I]I.

23

Q.0

•,,-" 315t-

,m

0U

I,.....1u')

0...

-.5

-.2

-.I

0 0

[] 0

57 0 0

106 z_ 0

159 r,, 0

.I0

/

/

f \

\ >--o_..

-Q,

/ \

\<

\

.. ½_x

/

.2 .4 .6

f

t I \\

,/

Pt, j/Pco 5

o...-o.

"3---,

\

\

j_s "_

.8 1.0 0 .2 .4

Frection of (]fterbody length, x/L

.6

.'b

Q\k

\

z_\

\

.8 1.0

(b) M = o.9o.

Figure 7.- Continued.

24

-.3

-.2

/

//

©

-.I

8, deg//

6 0 0 /r

O.

0

- /® 315 O 0"; U"

OU

1-

57 <> 0

Q.

106 a 0 //

.I0 .2

159 t,. 0

Jet off

_J

f___15 "/

/

\

\

\ _,,._ L

\,%

Pt, j/Pco : 5

/ \>..

\l/ - . r_'l

.." \ "o.,\

.- ,._

<:/ --,,, &\

q

\

\.4 .6 .8 1.0 0 .2 .4

Fraction of afterbod/ length, x/_,

\{,.

(c) M = o.9_.

\

"z

_E

.6 .8 1.0

Figure 7.- Continued.

25

8, deg

6

O.(..)

.5 315

,,g0

57(/1

13_

106

159

-.3

-.2

-.I

0 0

[] 0

t_ 0

.I0

/

1/

t

/

/

/

.2

\

/

f_

' Jet oft;

©4

\ LJ

Pt,j/Poo : 5

7

.4 .6 .8 1.0 0 .2 .4

Fraction of afterbody length, x//,

./ \/

(d) M = 1.O0.

.6

\

\

L

\

.8

\

\

\

\

I.O

Figure 7.- Continued.

26

8, deg

6 0

0

--" 315c

o

0_0o

= 57

n

106 z_

159 t,,

-.5

o 0

,,,,/v

0 0

/z

Jet off

i _ _ 3_ _ \

//

j/

/J

/

//

fJ_

ii

S ./

¥

\

Pt,'j/Poo - 5

I I I_-__

. __S _ "_

i

_x

t\/'

/

\L

v

\\

\"t

0 .2 .4 .6 .8 1.0 0 .2

Frocfion of ofterlI:ody length, x/7,

(e) M = 1.95.

.4 .6 .8 1,0

Figure 7-- Concluded.

27

-.6

--.4

-.2

0

.2

Jet off

M = 0.80

Pt, j/Po0 5

I

J 5,_

--o.

-.6

0

t-Q)

0

O_

--.4

0

.2

.4

--.6

--.4

--2

0

M =0.90

Config "ation

o I,/9= 16oLJ 12", /_eqv = 8°o nT, /_:8o

/

/

M =0.95

o"

/_ C-_ %

/o ,\

%

"-x

.20 .2 .4 .6 .8 1.0 0 .2 .4 .6 .8 1.0

Frocfion of efterbody length, x/l,

(a) M = 0.80_ 0.90, and 0.95.

Figure 8.- Comparisons of pressure distributions for the afterbody con-

figurations tested. 9 _ 0°.

28

E1

L)

C

0

:DU)oO

13-

-.6

-.4

--.2

0

.2

--.6

--.4

-.2

0

.20

Jet ot;f

/ \J

7

<i_'f_ _: ...... :y_--_ -_'° _

Configuration

o I, B=i6 °

1:3 Z, Beqv =8 °

<> m,/3 ::8o

M = 1.00-- Pt, j/Poo =5 _

I,.>

f..j_

l Jr

Y

_,.d

_--, d<

M = 1.05

.2 .4 .6 .8 1.0 0 .2 .4

Fraction of ofterbody length, x//,

.6 .8 1.0

(b) M = 1.00 and 3.05.

Figure 8.- Concluded.

29

C.m

b5

__ . c°_

i [] "6

__L_ _q}

--! i /_ rL

i i

C.i

(3, _ c

-') x

_- C0 C

_c_ _.__ 0 rn

II o oI

pi--

7

_ ,

__J ,,

, I ]

I I I I

d 0 iiueoy_eo3 eJnsseJd

_r

Q

i/

4._

0

q)

°_

0 (I.r4

.r4

•,-t e_

00

g-i

_o

4._ .,44._

og0

% oo

I1)

r/l .r-Im %_ °r-4

I I_ ,-4

0 .,--4

+_ I1)¢)-_

0 0

+_ _0 X_ -,-4

_ m

I

&

g

3o

o.

I

o

o-

0-

,i

/+/

J

i I

++

!

I/

/t

//++

JI

?]

j

In

.)II' Ii,/

]/

J_

iJ

i

q

oq'_ __

II

/u,3 !o.- tii

_ x

+i ¸

/1/

/!

I

/;

/

k

S

o

(F̧ o

o.+1

_E

0 '_" 0 ,_,IL"J ¢',.I

I

t

!

!/

I

\\\

I/]

/

//

o'O 0 'J,ua!o!jjaoo 5oJp-aJnssaJd ApoqJa,t,t V

/!P

Qoq q

1£3

o"+5L

OJ

icu- 6

o

Lc>

_t

r_

{-j

0

0

0o,-i4._

%

I

O4._

Q)°¢-._

4-_

_,-I F'4_ H©

O O

_ -r--t

m O

O

%I1)

O

O°,-t

°r-t%

>

!

c;

-r-t

31

0ii

E

O ....

Obd \ .'o..

)

,.-..

-%.%

_o

14-)

cy

0

ii

¢,J..,___0 ,,

>8o .o-- IT. N.*I- _

cc E2c_ o._Ooo _

"7 --_ _ _ 0 "_

...... C) N --

0 0 0 _ _- _._

I Io oO II

o.."1

L

\\

1

aO _,.q-L

i I__h _ L

co _ 0 (,D

_J_

oJ 00 _'- (

f

0 _ "I,7::_0

ii

o. _ c_-- I ed -- --

d '_-i0 ,_ua!3!ilao 3 isnJ41 alZZOU_,_Jow!Jd

00 q- O

_4

©

_nIn

I

O

4o

","D

4o.,-I

L0

14")

04

I

4-_

.,-I

_Y- tO

_. o +_

o,1 ! _o N_ s

0

%

%

0

o

4o

°_

I

,-4

t_

.52

u_"0

8

=_ --,4

t"

*$-_ 2.0

i_ 1.6

'M--O.e'O

/ ,/

1.6 .', /

// _,/

1.2

4l/

//.4 /

4//

,v

/o

1.2

.8

.4

o

--.4

M=i.oo

f

/

, //

,//

7//

/')

/r

/i I_ /

/

/

/

p/

I 2 3 4 5

-s F,7 :oo9-

,/,)5" ,_

,,/ ",/,/ .g

/,,', ,'i

M =).o5

M =0.95I

,/

,!/

/

,//" .,J

s) /

/

I 2 3 4 5

,..,/ 0 Wp,./ Tt.p--0

" 9 -- EjectorjetthrustcoefficientF

," _ ------Idealconvergent-nozzlejet

,_/ thrust coefficient, CFi,c

/

I 2 3 4 5

Jet total-pressure ratio, Pt,j/Poo

Figure 12.- Variation of ejector jet-thrust _oefficient with jet total-

pressure ratio for afterbody configuration llI. w s = O.

33

1.10

WS

Wp

Config. TIT

M = 0.90

0 0[] .0190 .059

Tt_ sTt_ p

Ref. 9

Ejector with diameter

ratio of 1.40 and

spacing ratio of 0.80:5

M=O

0 (Fig. 7(g) of

ref. 9)_' .059

u

ii0

°,,-_1

0

o"°_

o

-1IB

t-

I .00ii,

\ t\ <>

\I <>

.90 I

80

IJ

f

/[:]

,.z ..._)

f

.?O

.60I 2 3 4 5

Jet total pressure ratio, Pt, J/pooz

Figure 13.- Effect of secondary to primary weight flow ratio on jet

thrust ratio for afterbody configuration III.

6

34

O

o.

_'%L.

LO

re)

OJ

O o ii o

. C _,D >Og

.oT, g,,

E I I '

Lr) _- _

qO

----¢_o. I I\

__b

o

ii

\

I#)

re}

\

'r-.%1

m _t Q _o_ ea -

0

o

ii

-- I Od -- --

(]O -- dO '_.ua!o!llaoo BDJp-snu!LU-,SnJLi.I.

IC)

I"

#

Q.I

o_-_

o

r/l

ul

!

0

4._

°r-_

_d_ ,---t

• el I1)

°H

o_o o

!

4-_

0

o

°r'l

!

°_

55

20

1.8

Configuration

I, B = 16° Pt,j

Tr _eqv = 8° Po0

--TIT, /9: 8 °

5

4 //

//

580 90 I00

M

I I I

I10

r_0

ILI_

(,.)

c

._o

.D

(D00

0K-

"oIcn

c

EZ

L.

F--

,,., ....

12

I0

8

"x x,_

.6.80 .84 .88 .92 .96 1.00 104 1.08

Moch number, M

Figure 15.- Variation of thrust-minus-drag coefficient with Mach number

for the different afterbodies at a scheduled pressure ratio, w s = O.

36

Wp _/Tt, p

0 .04

0

[]

Config.

I, _ = 16°

Tr _eqv = B°

Tl'r /_=8 o

.5

.4

C30

t-q.)

._o .5

O.(.3

0

_o .2q3

UJ

.I

//

J_

Pt, j

PCO

/

/

4 ./

3.80 .90

//

/ t

f

f

/

/

M1.00

I

I/

I

/

J

I.I0

0.80 .84 88 92 .9G 1.00 1.04

Moch number, M

1.08

Figure 16.- Variation of effective drag coefficient with Mach number

for the different afterbodies at a sch_duled jet total-pressure

ratio.

3T

cO

I

OJ_

!

0 0

H _-_ .__°_

°__ °_

o o _o

_- 0 (J 0D C C C

.__- __

o r_

' III

//'

/,

/

,7

!

//

/

d v. xowVb

' xou.=v x 0-_-I

'o!_D_ oaJo x .l.ua!ojj.j.aoo 5DJp-snu!w-,tsnJq_L

Q0O

O

OO

coaO

cO

0o _.

O

,,r-t

_A

_o_ .r-4

.-I 4-_

O _°_-._

t__ m%

m

%

m %

O.-_ej

om

•,-4 _

O _

O!

r--I

°r--I

38

CO tO _ cJ

d V xo.w V bx

xouj v 0 -._-I

_O!$OJ OaJO X _uaj.0!jja00 6o_p-snu!w-isnJqi

0

-p

°.o

O

I.,_ 0

0.-_ __ o %

_ .,--I

O 0 _

I 4-_

._ _o °_

O _ _

I 0-o ,---t

_ or't

O o0

_ 0

• -0 0cO 0

%-,-_IlJ % +._

or'-t

._ • ,_-<_._ . .

Z

Q;

Z0

_ _u _ _".'_o =_ '_,_-_'_

_I ...... ._ _ =,_._ -,-'_ _

_,_ _ _1 _ _ _ 0 _